US8096133B2 - Method and apparatus for cooling and dilution tuning a gas turbine combustor liner and transition piece interface - Google Patents

Method and apparatus for cooling and dilution tuning a gas turbine combustor liner and transition piece interface Download PDF

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Publication number
US8096133B2
US8096133B2 US12/153,020 US15302008A US8096133B2 US 8096133 B2 US8096133 B2 US 8096133B2 US 15302008 A US15302008 A US 15302008A US 8096133 B2 US8096133 B2 US 8096133B2
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cooling
dilution
sleeve
liner
combustor liner
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US20090282833A1 (en
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William K. Hessler
Predrag Popovic
Charles Nyberg
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General Electric Co
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General Electric Co
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Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: HESSLER, WILLIAM, NYBERG, CHARLES, POPOVIC, PREDRAG
Priority to FR0953044A priority patent/FR2931225A1/en
Priority to JP2009115095A priority patent/JP5313038B2/en
Priority to DE102009025795A priority patent/DE102009025795A1/en
Priority to CNA200910145667XA priority patent/CN101581450A/en
Publication of US20090282833A1 publication Critical patent/US20090282833A1/en
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/26—Controlling the air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00012—Details of sealing devices
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03042—Film cooled combustion chamber walls or domes

Definitions

  • This invention relates to internal cooling within a gas turbine engine, and more particularly, to an apparatus and a method for providing better and more uniform cooling in a transition or interface region between a combustor liner and a transition piece.
  • Tuning of the combustor which can only be done after the turbine is operational, typically involves disassembly of the turbine and removal of the transition piece for drilling or welding dilution holes therein. This is a time-consuming and thus costly process.
  • the present invention relates to a combustor liner comprising a forward end and an aft end, the aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween; a plurality of cooling air entry holes formed in the cooling sleeve and a plurality of cooling air exit holes formed adjacent an aft edge of the liner such that, in use, cooling air flows through the cooling air entry holes and through the plenum, exiting the cooling air exit holes thereby cooling the aft end of the combustor liner, and affecting dilution tuning.
  • the invention in another exemplary but nonlimiting aspect, relates to a combustor liner comprising a liner forward end and a liner aft end, the liner aft end having a reduced diameter portion and a cooling sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween; a plurality of cooling air entry holes formed in the cooling sleeve and a plurality of cooling air exit holes formed adjacent an aft liner edge such that, in use, cooling air flows through the cooling air entry holes, and through the plenum, exiting the cooling air exit holes thereby cooling the aft liner end; wherein a compression seal is secured to an exterior surface of the cooling sleeve, directly radially outwardly of the plenum; and wherein the liner aft end includes an inwardly tapered portion leading to the reduced diameter portion, and an outwardly tapered portion leading to an annular collar, the cooling sleeve having a forward slee
  • the invention relates to a method of cooling an aft end of a combustor liner and associated annular seal comprising:
  • cooling air entry holes in an upstream end of the cooling sleeve and cooling air exit holes in the liner, proximate an aft edge thereof, such that, in use, cooling air flows through the cooling air entry holes into the plenum and through the cooling air exit holes.
  • FIG. 1 is a partial section illustrating a known gas turbine transition piece construction, with an associated combustor liner shown schematically;
  • FIG. 2 is an exploded partial perspective view of the aft end of a combustor liner in accordance with a known cooling configuration
  • FIG. 3 is a partial section through the aft end of a combustor liner in accordance with a first exemplary but nonlimiting embodiment of the invention
  • FIG. 4 is a partial section through the aft end of a combustor liner in accordance with a second exemplary but nonlimiting embodiment of the invention.
  • FIG. 5 is a partial section through the aft end of a combustor liner in accordance with a third exemplary but nonlimiting embodiment of the invention.
  • a typical gas turbine includes a transition piece 10 by which the hot combustion gases from an upstream combustor as represented by the schematically-illustrated combustor liner 12 are passed to the first stage of a turbine represented at 14 .
  • Flow from the gas turbine compressor exits an axial diffuser 16 and enters into a compressor discharge case 18 .
  • about 50% of the compressor discharge air passes through apertures 20 formed along and about a transition piece impingement sleeve 22 for flow in an annular region or annulus 24 (or, second flow annulus) between the transition piece 10 and the radially outer transition piece impingement sleeve 22 .
  • FIG. 2 a current cooling configuration for the aft end of liner 12 (similar to the arrangement disclosed in U.S. Pat. No. 7,010,921) is shown to include an associated compression-type seal 26 , commonly referred to as a “hula seal”, mounted on the cover plate 28 , adjacent a tapered transition region 30 of the liner.
  • a compression-type seal 26 commonly referred to as a “hula seal”
  • FIG. 2 shows only “slices” of the otherwise annular component parts.
  • the cylindrical cover plate 28 is mounted on the liner to form a mounting surface for the annular compression seal 26 and to form a portion of the axial airflow channels 34 .
  • liner 12 has a plurality of axial raised sections or ribs 32 which extend over a portion of aft end of the liner 12 .
  • the cover plate 28 and ribs 32 together define the respective airflow channels 34 . These channels are substantially parallel and extend over a portion of aft end of liner 12 .
  • Cooling air is introduced into the channels 34 through an annular air inlet slot or opening 36 (or holes 36 A) at the forward end of the channels. The cooling air flows into and through the channels 34 and exits through openings 38 at an aft end of the liner.
  • FIG. 3 one exemplary but nonlimiting embodiment of the invention shows the aft end of the liner 40 formed with a radially inward taper 42 , establishing a reduced diameter portion 44 at the downstream end of the liner:
  • a radially outwardly tapered portion 46 expands the liner to a larger diameter, with an annular collar 47 terminating at an aft edge 48 .
  • a tubular, cylindrical cooling sleeve 50 extends from a point about midway along the tapered portion 42 of the liner, to an aft edge 54 adjacent the terminal edge 48 of the liner.
  • the sleeve 50 may be welded to the collar portion 47 of the liner.
  • a spring seal (or hula seal) 56 is secured to the aft edge of the cooling sleeve 50 (by, for example, welding) and enables sealing engagement with the interior surface of the transition piece (not shown).
  • Cooling air entry holes 58 are provided in an annular array about the cooling sleeve 50 , at a location proximate the forward edge 52 of the sleeve. Thus, cooling air flowing through the air entry holes 58 , enters a cooling plenum 60 between the reduced diameter portion 44 of the liner and the cooling sleeve 50 . Air flowing through the plenum exits through an annular array of cooling air exit holes 62 formed in the outwardly tapered portion 46 of the liner.
  • section of the liner including the inwardly tapered portion 42 , the reduced diameter portion 44 , outwardly tapered portion 46 and collar 47 may be separately formed and welded to the liner at a location indicated at 63 , for example.
  • a relatively tight fitting collar 51 could be applied over the sleeve 50 , axially behind the seal 56 .
  • the collar 51 could have a series of circumferentially-spaced holes 59 in selected locations such that the collar could then be rotated to place some or all of the holes 59 into partial or full alignment with holes 58 to thereby achieve the desired cooling and tuning dilution characteristics without having to remove the transition piece and add holes to the sleeve 50 .
  • FIG. 4 illustrates another exemplary but nonlimiting embodiment of the invention.
  • This embodiment represents a modification of the known configuration shown in FIG. 2 .
  • a new aft end piece 64 is added to the liner, and the original transition region ( 30 in FIG. 2 ) now serves as a cooling sleeve 76 , establishing a plenum 66 therebetween.
  • the aft end piece 64 is similar to the aft end in FIG. 3 , with an inwardly tapered portion 68 leading to a reduced diameter portion 70 .
  • An outwardly tapered aft end portion 72 leads to a terminal collar and edge (not shown, but similar to FIG. 3 ).
  • Cooling an exit holes 74 are formed in an annular array, about the tapered end portion 72 .
  • the cooling sleeve 76 also includes an inwardly tapered portion 78 leading to a reduced diameter portion 80 terminating at an aft end terminal edge (not shown).
  • the cooling sleeve is similar to the sleeve shown in FIG. 2 , but here, the channels are closed by eliminating the forward opening ( 36 in FIG. 2 ).
  • the cover plate 82 and reduced diameter portion 80 could be made as a single piece, eliminating the channels 84 .
  • FIG. 5 another exemplary and preferred embodiment of the invention is illustrated, which represents a simplified version of the cooling arrangement in FIG. 4 .
  • the aft end piece 88 of the liner 90 is again formed with a inwardly tapered portion 92 , a reduced diameter portion 94 , an outwardly tapered aft end portion 96 , and a flat terminal ring or collar 98 defining the aft edge 100 of the liner.
  • a cooling sleeve 102 is welded to the ring or collar 98 , of the liner, extending away from the terminal edge 100 , and including a tapered portion 104 merging with a forward edge portion 106 where a slip joint is formed with the aft end piece 88 , adjacent a forward edge 108 where the aft end piece 88 is joined to the liner 90 .
  • This arrangement creates a plenum 110 between the aft end piece 88 and the cooling sleeve 102 .
  • Cooling air entry holes 112 in the sleeve 102 permit cooling and dilution air to enter the plenum 110 , and to exit via holes 114 formed in the outwardly tapered portion 96 of the aft end piece 88 .
  • a spring finger or hula seal 116 is welded to the aft end of the cooling sleeve 102 .
  • FIGS. 3-5 dumps cooling air under the hula seal, cooling the seal along with the aft end of the combustor liner.
  • tuning of the combustor liner as it relates to the optimization of dilution air can be accomplished without having to remove the transition piece, by modifying or adding cooling air entry and/or end holes to the cooling sleeve and/or liner.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Gas Burners (AREA)
  • Spray-Type Burners (AREA)

Abstract

A combustor liner includes a forward end and an aft end, the aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween. A plurality of cooling and dilution air entry holes are formed in the cooling and dilution sleeve and a plurality of cooling and dilution air exit holes formed adjacent an aft edge of the liner such that, in use, cooling and dilution air flows through the cooling and dilution air entry holes, and through the plenum, exiting the cooling and dilution air exit holes, thereby cooling and dilution tuning the aft end of the combustor liner without having to remove the transition piece.

Description

BACKGROUND OF THE INVENTION
This invention relates to internal cooling within a gas turbine engine, and more particularly, to an apparatus and a method for providing better and more uniform cooling in a transition or interface region between a combustor liner and a transition piece.
Traditional gas turbine combustors use diffusion (i.e., non-premixed) combustion in which fuel and air enter the combustion chamber separately. The process of mixing and burning produces flame temperatures exceeding 3900° F. Since conventional combustors and/or transition pieces are generally capable of withstanding a maximum temperature on the order of only about 1500° F. for about ten thousand hours (10,000 hrs.), steps to protect the combustor and/or transition piece must be taken. This has typically been done by film-cooling which involves introducing relatively cool compressor air into a plenum formed by an impingement cooling sleeve surrounding the transition piece and a flow sleeve surrounding the combustor liner. This cooling air is ultimately reverse-flowed into the combustor where it mixes with fuel for combustion and dilution tuning.
Various techniques have been employed to cool the aft end of the combustor liner (that end adjacent the transition piece) and the compression seal (or “hula” seal) typically used at the interface of the transition piece and combustor liner. See, for example, U.S. Pat. No. 6,098,397 which discloses providing an array of concavities on the outside surface of the liner to enhance heat transfer. Another technique is disclosed in U.S. Pat. No. 7,010,921 where the aft end of the combustor liner is provided with a plurality of axially extending ribs or turbulators about its circumference, covered with a sleeve or cover plate, thus forming a series of cooling channels. Cooling air is introduced into the channels through air inlet slots or openings at the forward end of the channels, and exits into the transition piece which is telescoped over the aft end of the liner.
Tuning of the combustor (including the cooling configuration), which can only be done after the turbine is operational, typically involves disassembly of the turbine and removal of the transition piece for drilling or welding dilution holes therein. This is a time-consuming and thus costly process.
There remains a need, therefore, for a cooling arrangement that provides effective, uniform cooling of the aft end of the combustor liner/transition piece interface, but that also simplifies the combustor tuning process.
BRIEF DESCRIPTION OF THE INVENTION
In one exemplary but nonlimiting aspect, the present invention relates to a combustor liner comprising a forward end and an aft end, the aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween; a plurality of cooling air entry holes formed in the cooling sleeve and a plurality of cooling air exit holes formed adjacent an aft edge of the liner such that, in use, cooling air flows through the cooling air entry holes and through the plenum, exiting the cooling air exit holes thereby cooling the aft end of the combustor liner, and affecting dilution tuning.
In another exemplary but nonlimiting aspect, the invention relates to a combustor liner comprising a liner forward end and a liner aft end, the liner aft end having a reduced diameter portion and a cooling sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween; a plurality of cooling air entry holes formed in the cooling sleeve and a plurality of cooling air exit holes formed adjacent an aft liner edge such that, in use, cooling air flows through the cooling air entry holes, and through the plenum, exiting the cooling air exit holes thereby cooling the aft liner end; wherein a compression seal is secured to an exterior surface of the cooling sleeve, directly radially outwardly of the plenum; and wherein the liner aft end includes an inwardly tapered portion leading to the reduced diameter portion, and an outwardly tapered portion leading to an annular collar, the cooling sleeve having a forward sleeve end engaged with the liner at a location upstream of the inwardly tapered portion, and an aft sleeve end fixed to the collar.
In still another exemplary but non limiting aspect, the invention relates to a method of cooling an aft end of a combustor liner and associated annular seal comprising:
forming an aft end portion of the liner with a reduced diameter portion;
locating a cooling sleeve about the reduced diameter portion, in radially spaced relationship thereto so as to create an annular plenum;
forming cooling air entry holes in an upstream end of the cooling sleeve and cooling air exit holes in the liner, proximate an aft edge thereof, such that, in use, cooling air flows through the cooling air entry holes into the plenum and through the cooling air exit holes.
The invention will now be described in detail in connection with the drawings identified below.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a partial section illustrating a known gas turbine transition piece construction, with an associated combustor liner shown schematically;
FIG. 2 is an exploded partial perspective view of the aft end of a combustor liner in accordance with a known cooling configuration;
FIG. 3 is a partial section through the aft end of a combustor liner in accordance with a first exemplary but nonlimiting embodiment of the invention;
FIG. 4 is a partial section through the aft end of a combustor liner in accordance with a second exemplary but nonlimiting embodiment of the invention; and
FIG. 5 is a partial section through the aft end of a combustor liner in accordance with a third exemplary but nonlimiting embodiment of the invention.
 DETAILED DESCRIPTION OF THE DRAWINGS
With reference to FIG. 1, a typical gas turbine includes a transition piece 10 by which the hot combustion gases from an upstream combustor as represented by the schematically-illustrated combustor liner 12 are passed to the first stage of a turbine represented at 14. Flow from the gas turbine compressor exits an axial diffuser 16 and enters into a compressor discharge case 18. In some transition piece designs, about 50% of the compressor discharge air passes through apertures 20 formed along and about a transition piece impingement sleeve 22 for flow in an annular region or annulus 24 (or, second flow annulus) between the transition piece 10 and the radially outer transition piece impingement sleeve 22. The remaining approximately 50% of the compressor discharge flow passes into holes in an upstream combustion liner flow sleeve (not shown) and into an annulus between the flow sleeve and the liner 12 where it mixes with the air in annulus 24 before being reverse-flowed into the combustion chamber.
A known arrangement for coupling of the transition piece and impingement sleeve with the combustor liner and flow sleeve is disclosed in, for example, U.S. Pat. No. 7,010,921, and need not be described in further detail here.
In FIG. 2, a current cooling configuration for the aft end of liner 12 (similar to the arrangement disclosed in U.S. Pat. No. 7,010,921) is shown to include an associated compression-type seal 26, commonly referred to as a “hula seal”, mounted on the cover plate 28, adjacent a tapered transition region 30 of the liner. It will be appreciated that FIG. 2 shows only “slices” of the otherwise annular component parts. Thus, the cylindrical cover plate 28 is mounted on the liner to form a mounting surface for the annular compression seal 26 and to form a portion of the axial airflow channels 34. More specifically, liner 12 has a plurality of axial raised sections or ribs 32 which extend over a portion of aft end of the liner 12. The cover plate 28 and ribs 32 together define the respective airflow channels 34. These channels are substantially parallel and extend over a portion of aft end of liner 12. Cooling air is introduced into the channels 34 through an annular air inlet slot or opening 36 (or holes 36A) at the forward end of the channels. The cooling air flows into and through the channels 34 and exits through openings 38 at an aft end of the liner.
Turning now to FIG. 3, one exemplary but nonlimiting embodiment of the invention shows the aft end of the liner 40 formed with a radially inward taper 42, establishing a reduced diameter portion 44 at the downstream end of the liner: A radially outwardly tapered portion 46 expands the liner to a larger diameter, with an annular collar 47 terminating at an aft edge 48. A tubular, cylindrical cooling sleeve 50 extends from a point about midway along the tapered portion 42 of the liner, to an aft edge 54 adjacent the terminal edge 48 of the liner. The sleeve 50 may be welded to the collar portion 47 of the liner. A spring seal (or hula seal) 56 is secured to the aft edge of the cooling sleeve 50 (by, for example, welding) and enables sealing engagement with the interior surface of the transition piece (not shown).
Cooling air entry holes 58 are provided in an annular array about the cooling sleeve 50, at a location proximate the forward edge 52 of the sleeve. Thus, cooling air flowing through the air entry holes 58, enters a cooling plenum 60 between the reduced diameter portion 44 of the liner and the cooling sleeve 50. Air flowing through the plenum exits through an annular array of cooling air exit holes 62 formed in the outwardly tapered portion 46 of the liner.
Note that the section of the liner including the inwardly tapered portion 42, the reduced diameter portion 44, outwardly tapered portion 46 and collar 47 may be separately formed and welded to the liner at a location indicated at 63, for example. In order to tune the combustor with this arrangement, it is only necessary to drill additional cooling air entry holes in the sleeve 50 as needed, without also having to remove the transition piece. This is a tune-saving design.
In a variation of this design, a relatively tight fitting collar 51 could be applied over the sleeve 50, axially behind the seal 56. The collar 51 could have a series of circumferentially-spaced holes 59 in selected locations such that the collar could then be rotated to place some or all of the holes 59 into partial or full alignment with holes 58 to thereby achieve the desired cooling and tuning dilution characteristics without having to remove the transition piece and add holes to the sleeve 50.
FIG. 4 illustrates another exemplary but nonlimiting embodiment of the invention. This embodiment represents a modification of the known configuration shown in FIG. 2. In this case, however, a new aft end piece 64 is added to the liner, and the original transition region (30 in FIG. 2) now serves as a cooling sleeve 76, establishing a plenum 66 therebetween. The aft end piece 64 is similar to the aft end in FIG. 3, with an inwardly tapered portion 68 leading to a reduced diameter portion 70. An outwardly tapered aft end portion 72 leads to a terminal collar and edge (not shown, but similar to FIG. 3). Cooling an exit holes 74 are formed in an annular array, about the tapered end portion 72. The cooling sleeve 76 also includes an inwardly tapered portion 78 leading to a reduced diameter portion 80 terminating at an aft end terminal edge (not shown). The cooling sleeve is similar to the sleeve shown in FIG. 2, but here, the channels are closed by eliminating the forward opening (36 in FIG. 2). In fact, the cover plate 82 and reduced diameter portion 80 could be made as a single piece, eliminating the channels 84.
Cooling air now enters the cooling air entry holes 86 formed about the tapered portion 78, flows through the plenum 66 and exits through cooling air exit holes 74 formed in the tapered aft end portion 72.
Turning now to FIG. 5, another exemplary and preferred embodiment of the invention is illustrated, which represents a simplified version of the cooling arrangement in FIG. 4. Specifically, the aft end piece 88 of the liner 90 is again formed with a inwardly tapered portion 92, a reduced diameter portion 94, an outwardly tapered aft end portion 96, and a flat terminal ring or collar 98 defining the aft edge 100 of the liner. A cooling sleeve 102 is welded to the ring or collar 98, of the liner, extending away from the terminal edge 100, and including a tapered portion 104 merging with a forward edge portion 106 where a slip joint is formed with the aft end piece 88, adjacent a forward edge 108 where the aft end piece 88 is joined to the liner 90. This arrangement creates a plenum 110 between the aft end piece 88 and the cooling sleeve 102. Cooling air entry holes 112 in the sleeve 102 permit cooling and dilution air to enter the plenum 110, and to exit via holes 114 formed in the outwardly tapered portion 96 of the aft end piece 88. A spring finger or hula seal 116 is welded to the aft end of the cooling sleeve 102.
The above-described cooling and dilution arrangements illustrated in FIGS. 3-5 dumps cooling air under the hula seal, cooling the seal along with the aft end of the combustor liner. At the same time, tuning of the combustor liner as it relates to the optimization of dilution air can be accomplished without having to remove the transition piece, by modifying or adding cooling air entry and/or end holes to the cooling sleeve and/or liner.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.

Claims (18)

1. A combustor liner comprising a forward end and an aft end, the aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween; a plurality of cooling and dilution air entry holes formed in said cooling sleeve and a plurality of cooling and dilution air exit holes formed in said combustor liner adjacent an aft edge of said combustor liner such that, in use, cooling air flows through said cooling and dilution air entry holes and through said plenum, exiting said plenum via said cooling and dilution air exit holes thereby cooling said aft end of said combustor liner.
2. The combustor liner of claim 1 wherein a compression seal is secured to an exterior surface of said cooling and dilution sleeve, directly radially outwardly of said plenum.
3. The combustor liner of claim 1 wherein said aft end includes an inwardly tapered portion leading to said reduced diameter portion, and an outwardly tapered portion leading to an annular collar, and further wherein a forward edge of said cooling and dilution sleeve engages said inwardly tapered portion and an aft edge of said cooling and dilution sleeve is secured to said collar.
4. The combustor liner of claim 1 wherein said aft end includes an inwardly tapered portion leading to said reduced diameter portion, and an outwardly tapered portion leading to an annular collar, and further wherein said cooling and dilution sleeve has a forward sleeve end engaged with the liner at a location upstream of said inwardly tapered portion and an aft sleeve end fixed to said collar.
5. The combustor liner of claim 3 wherein said cooling and dilution air entry holes are formed in an annular array adjacent said forward edge of said cooling and dilution sleeve.
6. The combustor liner of claim 5 wherein said cooling and dilution air exit holes are formed in an annular array in said outwardly tapered portion of said combustor liner.
7. The combustor liner of claim 4 wherein said cooling and dilution sleeve has an inwardly tapered portion proximate a forward edge thereof, and further wherein said cooling and dilution air entry holes are formed in an annular array in said inwardly tapered portion of said cooling and dilution sleeve, and said cooling and dilution air exit holes are formed in said outwardly tapered portion of said combustor liner.
8. The combustor of claim 7 wherein a compression seal is secured to an exterior surface of said cooling and dilution sleeve, directly radially outwardly of said plenum.
9. The combustor liner of claim 7 wherein said forward end of said cooling sleeve forms a slip joint with said combustor liner.
10. A combustor liner comprising a liner forward end and a liner aft end, the liner aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween; a plurality of cooling and dilution air entry holes formed in said cooling and dilution sleeve and a plurality of cooling and dilution air exit holes formed adjacent an aft liner edge such that, in use, cooling air flows through said cooling and dilution air entry holes, and through said plenum, exiting said cooling and dilution air exit holes thereby cooling said aft liner end;
wherein a compression seal is secured to an exterior surface of said cooling and dilution sleeve, directly radially outwardly of said plenum; and
wherein said combustor liner aft end includes an inwardly tapered portion leading to said reduced diameter portion, and an outwardly tapered portion leading to an annular collar, said cooling and dilution sleeve having a forward sleeve end engaged with the liner at a location upstream of said inwardly tapered portion, and an aft sleeve end fixed to said collar.
11. The combustor liner of claim 10 wherein said cooling and dilution air exit holes are located in said outwardly tapered portion of said combustor liner.
12. The combustor liner of claim 10 wherein said cooling and dilution sleeve has an inwardly tapered portion proximate a forward edge thereof, and further wherein said cooling and dilution air entry holes are formed in an annular array in said inwardly tapered portion of said cooling and dilution sleeve, and said cooling and dilution air exit holes are formed in said outwardly tapered portion of said liner.
13. The combustor liner of claim 10 wherein said forward end of said cooling and dilution sleeve forms a slip joint with said combustor liner.
14. A method of cooling and dilution tuning an aft end of a combustor liner and associated annular seal comprising:
forming an aft end portion of the liner with a reduced diameter portion;
locating a cooling and dilution sleeve about said reduced diameter portion, in radially spaced relationship thereto so as to create an annular plenum;
forming cooling and dilution air entry holes in an upstream end of said cooling and dilution sleeve and cooling and dilution air exit holes in said combustor liner, proximate an aft edge thereof, such that, in use, cooling air flows through said cooling and dilution air entry holes into said plenum and through said cooling and dilution air exit holes.
15. The method of claim 14 wherein said cooling and dilution air entry holes are formed in an inwardly tapered transition region of said liner that leads to said reduced diameter portion.
16. The method of claim 15 wherein said cooling and dilution air exit holes are located in an outwardly tapered surface of said combustor liner, leading to said aft edge.
17. The method of claim 14 wherein a compression seal is supported on said cooling and dilution sleeve radially outwardly of said plenum.
18. The method of claim 14 including locating a rotatable collar in close-fitting relationship about a portion of said cooling and dilution sleeve said collar having holes therein that are selectively alignable with said cooling and dilution air entry holes to thereby achieve desired cooling and dilution tuning characteristics.
US12/153,020 2008-05-13 2008-05-13 Method and apparatus for cooling and dilution tuning a gas turbine combustor liner and transition piece interface Expired - Fee Related US8096133B2 (en)

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US12/153,020 US8096133B2 (en) 2008-05-13 2008-05-13 Method and apparatus for cooling and dilution tuning a gas turbine combustor liner and transition piece interface
FR0953044A FR2931225A1 (en) 2008-05-13 2009-05-07 METHOD FOR CONTROLLING COOLING AND DILUTION AT A COMBUSTION CHAMBER INTERFACE AND A GAS COMBUSTION CHAMBER
JP2009115095A JP5313038B2 (en) 2008-05-13 2009-05-12 Method and apparatus for cooling and dilution adjustment of a gas turbine combustor liner and transition piece joint
DE102009025795A DE102009025795A1 (en) 2008-05-13 2009-05-13 A method and apparatus for cooling and blending a junction between a gas turbine combustor flame tube and a transition piece
CNA200910145667XA CN101581450A (en) 2008-05-13 2009-05-13 Method and apparatus for cooling and dilution tuning gas turbine combustor linear and transition piece interface

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Citations (38)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3777484A (en) * 1971-12-08 1973-12-11 Gen Electric Shrouded combustion liner
US4413477A (en) * 1980-12-29 1983-11-08 General Electric Company Liner assembly for gas turbine combustor
US4628694A (en) * 1983-12-19 1986-12-16 General Electric Company Fabricated liner article and method
US4719748A (en) * 1985-05-14 1988-01-19 General Electric Company Impingement cooled transition duct
US4872312A (en) * 1986-03-20 1989-10-10 Hitachi, Ltd. Gas turbine combustion apparatus
US4903477A (en) * 1987-04-01 1990-02-27 Westinghouse Electric Corp. Gas turbine combustor transition duct forced convection cooling
US4916906A (en) * 1988-03-25 1990-04-17 General Electric Company Breach-cooled structure
US5127221A (en) * 1990-05-03 1992-07-07 General Electric Company Transpiration cooled throat section for low nox combustor and related process
US5285631A (en) * 1990-02-05 1994-02-15 General Electric Company Low NOx emission in gas turbine system
US5454221A (en) * 1994-03-14 1995-10-03 General Electric Company Dilution flow sleeve for reducing emissions in a gas turbine combustor
US5735126A (en) * 1995-06-02 1998-04-07 Asea Brown Boveri Ag Combustion chamber
US5784876A (en) * 1995-03-14 1998-07-28 European Gas Turbines Limited Combuster and operating method for gas-or liquid-fuelled turbine arrangement
US6098397A (en) * 1998-06-08 2000-08-08 Caterpillar Inc. Combustor for a low-emissions gas turbine engine
US6134877A (en) * 1997-08-05 2000-10-24 European Gas Turbines Limited Combustor for gas-or liquid-fuelled turbine
US6334310B1 (en) 2000-06-02 2002-01-01 General Electric Company Fracture resistant support structure for a hula seal in a turbine combustor and related method
US6427446B1 (en) * 2000-09-19 2002-08-06 Power Systems Mfg., Llc Low NOx emission combustion liner with circumferentially angled film cooling holes
US6446438B1 (en) * 2000-06-28 2002-09-10 Power Systems Mfg., Llc Combustion chamber/venturi cooling for a low NOx emission combustor
US6484505B1 (en) * 2000-02-25 2002-11-26 General Electric Company Combustor liner cooling thimbles and related method
US6494044B1 (en) * 1999-11-19 2002-12-17 General Electric Company Aerodynamic devices for enhancing sidepanel cooling on an impingement cooled transition duct and related method
US6832482B2 (en) * 2002-06-25 2004-12-21 Power Systems Mfg, Llc Pressure ram device on a gas turbine combustor
US6865892B2 (en) * 2002-12-17 2005-03-15 Power Systems Mfg, Llc Combustion chamber/venturi configuration and assembly method
US20050132708A1 (en) 2003-12-22 2005-06-23 Martling Vincent C. Cooling and sealing design for a gas turbine combustion system
US20050144953A1 (en) * 2003-12-24 2005-07-07 Martling Vincent C. Flow sleeve for a law NOx combustor
US6928822B2 (en) 2002-05-28 2005-08-16 Lytesyde, Llc Turbine engine apparatus and method
US6935116B2 (en) * 2003-04-28 2005-08-30 Power Systems Mfg., Llc Flamesheet combustor
US6951109B2 (en) * 2004-01-06 2005-10-04 General Electric Company Apparatus and methods for minimizing and/or eliminating dilution air leakage in a combustion liner assembly
US6968693B2 (en) 2003-09-22 2005-11-29 General Electric Company Method and apparatus for reducing gas turbine engine emissions
US7082766B1 (en) 2005-03-02 2006-08-01 General Electric Company One-piece can combustor
US20060168967A1 (en) * 2005-01-31 2006-08-03 General Electric Company Inboard radial dump venturi for combustion chamber of a gas turbine
US7284378B2 (en) 2004-06-04 2007-10-23 General Electric Company Methods and apparatus for low emission gas turbine energy generation
US7299618B2 (en) 2003-12-09 2007-11-27 Mitsubishi Heavy Industries, Ltd. Gas turbine combustion apparatus
US20080092547A1 (en) * 2006-09-21 2008-04-24 Lockyer John F Combustor assembly for gas turbine engine
US7493767B2 (en) * 2004-06-01 2009-02-24 General Electric Company Method and apparatus for cooling combustor liner and transition piece of a gas turbine
US7524167B2 (en) * 2006-05-04 2009-04-28 Siemens Energy, Inc. Combustor spring clip seal system
US7631504B2 (en) * 2006-02-21 2009-12-15 General Electric Company Methods and apparatus for assembling gas turbine engines
US20100043441A1 (en) * 2008-08-25 2010-02-25 William Kirk Hessler Method and apparatus for assembling gas turbine engines
US7707835B2 (en) * 2005-06-15 2010-05-04 General Electric Company Axial flow sleeve for a turbine combustor and methods of introducing flow sleeve air
US7707836B1 (en) * 2009-01-21 2010-05-04 Gas Turbine Efficiency Sweden Ab Venturi cooling system

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5960127A (en) * 1982-09-29 1984-04-06 Toshiba Corp Combustor for gas turbine
US5117636A (en) * 1990-02-05 1992-06-02 General Electric Company Low nox emission in gas turbine system
JPH08285284A (en) * 1995-04-10 1996-11-01 Toshiba Corp Combustor structure for gas turbine
JP4709433B2 (en) * 2001-06-29 2011-06-22 三菱重工業株式会社 Gas turbine combustor
US7051532B2 (en) * 2003-10-17 2006-05-30 General Electric Company Methods and apparatus for film cooling gas turbine engine combustors
US7000396B1 (en) * 2004-09-02 2006-02-21 General Electric Company Concentric fixed dilution and variable bypass air injection for a combustor
US7571611B2 (en) * 2006-04-24 2009-08-11 General Electric Company Methods and system for reducing pressure losses in gas turbine engines

Patent Citations (42)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3777484A (en) * 1971-12-08 1973-12-11 Gen Electric Shrouded combustion liner
US4413477A (en) * 1980-12-29 1983-11-08 General Electric Company Liner assembly for gas turbine combustor
US4628694A (en) * 1983-12-19 1986-12-16 General Electric Company Fabricated liner article and method
US4719748A (en) * 1985-05-14 1988-01-19 General Electric Company Impingement cooled transition duct
US4872312A (en) * 1986-03-20 1989-10-10 Hitachi, Ltd. Gas turbine combustion apparatus
US4903477A (en) * 1987-04-01 1990-02-27 Westinghouse Electric Corp. Gas turbine combustor transition duct forced convection cooling
US4916906A (en) * 1988-03-25 1990-04-17 General Electric Company Breach-cooled structure
US5285631A (en) * 1990-02-05 1994-02-15 General Electric Company Low NOx emission in gas turbine system
US5127221A (en) * 1990-05-03 1992-07-07 General Electric Company Transpiration cooled throat section for low nox combustor and related process
US5454221A (en) * 1994-03-14 1995-10-03 General Electric Company Dilution flow sleeve for reducing emissions in a gas turbine combustor
US5784876A (en) * 1995-03-14 1998-07-28 European Gas Turbines Limited Combuster and operating method for gas-or liquid-fuelled turbine arrangement
US5735126A (en) * 1995-06-02 1998-04-07 Asea Brown Boveri Ag Combustion chamber
US6134877A (en) * 1997-08-05 2000-10-24 European Gas Turbines Limited Combustor for gas-or liquid-fuelled turbine
US6098397A (en) * 1998-06-08 2000-08-08 Caterpillar Inc. Combustor for a low-emissions gas turbine engine
US6494044B1 (en) * 1999-11-19 2002-12-17 General Electric Company Aerodynamic devices for enhancing sidepanel cooling on an impingement cooled transition duct and related method
US6484505B1 (en) * 2000-02-25 2002-11-26 General Electric Company Combustor liner cooling thimbles and related method
US6334310B1 (en) 2000-06-02 2002-01-01 General Electric Company Fracture resistant support structure for a hula seal in a turbine combustor and related method
US6446438B1 (en) * 2000-06-28 2002-09-10 Power Systems Mfg., Llc Combustion chamber/venturi cooling for a low NOx emission combustor
US6427446B1 (en) * 2000-09-19 2002-08-06 Power Systems Mfg., Llc Low NOx emission combustion liner with circumferentially angled film cooling holes
US6928822B2 (en) 2002-05-28 2005-08-16 Lytesyde, Llc Turbine engine apparatus and method
US6832482B2 (en) * 2002-06-25 2004-12-21 Power Systems Mfg, Llc Pressure ram device on a gas turbine combustor
US6865892B2 (en) * 2002-12-17 2005-03-15 Power Systems Mfg, Llc Combustion chamber/venturi configuration and assembly method
US6935116B2 (en) * 2003-04-28 2005-08-30 Power Systems Mfg., Llc Flamesheet combustor
US7260935B2 (en) 2003-09-22 2007-08-28 General Electric Company Method and apparatus for reducing gas turbine engine emissions
US6968693B2 (en) 2003-09-22 2005-11-29 General Electric Company Method and apparatus for reducing gas turbine engine emissions
US7299618B2 (en) 2003-12-09 2007-11-27 Mitsubishi Heavy Industries, Ltd. Gas turbine combustion apparatus
US20050132708A1 (en) 2003-12-22 2005-06-23 Martling Vincent C. Cooling and sealing design for a gas turbine combustion system
US7082770B2 (en) * 2003-12-24 2006-08-01 Martling Vincent C Flow sleeve for a low NOx combustor
US20050144953A1 (en) * 2003-12-24 2005-07-07 Martling Vincent C. Flow sleeve for a law NOx combustor
US6951109B2 (en) * 2004-01-06 2005-10-04 General Electric Company Apparatus and methods for minimizing and/or eliminating dilution air leakage in a combustion liner assembly
US7493767B2 (en) * 2004-06-01 2009-02-24 General Electric Company Method and apparatus for cooling combustor liner and transition piece of a gas turbine
US7284378B2 (en) 2004-06-04 2007-10-23 General Electric Company Methods and apparatus for low emission gas turbine energy generation
US7389643B2 (en) * 2005-01-31 2008-06-24 General Electric Company Inboard radial dump venturi for combustion chamber of a gas turbine
US20060168967A1 (en) * 2005-01-31 2006-08-03 General Electric Company Inboard radial dump venturi for combustion chamber of a gas turbine
US7082766B1 (en) 2005-03-02 2006-08-01 General Electric Company One-piece can combustor
US7707835B2 (en) * 2005-06-15 2010-05-04 General Electric Company Axial flow sleeve for a turbine combustor and methods of introducing flow sleeve air
US7631504B2 (en) * 2006-02-21 2009-12-15 General Electric Company Methods and apparatus for assembling gas turbine engines
US7524167B2 (en) * 2006-05-04 2009-04-28 Siemens Energy, Inc. Combustor spring clip seal system
US20080092547A1 (en) * 2006-09-21 2008-04-24 Lockyer John F Combustor assembly for gas turbine engine
US20100043441A1 (en) * 2008-08-25 2010-02-25 William Kirk Hessler Method and apparatus for assembling gas turbine engines
US7707836B1 (en) * 2009-01-21 2010-05-04 Gas Turbine Efficiency Sweden Ab Venturi cooling system
US7712314B1 (en) * 2009-01-21 2010-05-11 Gas Turbine Efficiency Sweden Ab Venturi cooling system

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120304657A1 (en) * 2011-06-06 2012-12-06 General Electric Company Lock leaf hula seal
US9328665B2 (en) 2012-08-03 2016-05-03 Rolls-Royce Deutschland Ltd & Co Kg Gas-turbine combustion chamber with mixing air orifices and chutes in modular design
US9163837B2 (en) 2013-02-27 2015-10-20 Siemens Aktiengesellschaft Flow conditioner in a combustor of a gas turbine engine
US10533750B2 (en) 2014-09-05 2020-01-14 Siemens Aktiengesellschaft Cross ignition flame duct

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US20090282833A1 (en) 2009-11-19
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